A cooling arrangement is provided for a turbine disc of a gas turbine engine. The turbine disc has a plurality of circumferentially spaced disc posts forming fixtures therebetween for a row of turbine blades. Each turbine blade has an attachment formation which engages at a respective fixture, a platform radially outwardly of the attachment formation such that the adjacent platforms of the row form an inner endwall for the working gas annulus of the engine, and an aerofoil which extends radially outwardly from the platform. A respective cavity is formed between an exposed radially outer surface of each disc post and the inner endwall. The cooling arrangement has at each disc post, a cooling plate located in the respective cavity and spaced radially outwardly from the exposed outer surface of the disc post to form a cooling channel between the cooling plate and the exposed outer surface.
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1. A cooling arrangement for a turbine disc of a gas turbine engine, the turbine disc having a plurality of circumferentially spaced disc posts forming fixtures therebetween for a row of turbine blades, each turbine blade having an attachment formation which engages at a respective fixture, a platform radially outwardly of the attachment formation such that the adjacent platforms of the row form an inner endwall for a working gas annulus of the engine, and an aerofoil which extends radially outwardly from the platform, wherein a respective cavity is formed between an exposed radially outer surface of each disc post and the inner endwall;
wherein the cooling arrangement comprises:
at each disc post, a cooling plate having one or more exit holes formed therein for transferring spent coolant from a cooling channel to the cavity, each cooling plate being removably locatable in the respective cavity and spaced radially outwardly from the exposed outer surface of the disc post and spaced radially inwardly from the platform so that the entirety of the cooling plate does not touch the platform and forms the cooling channel between the cooling plate and the exposed outer surface, each cooling plate covering substantially the entire exposed outer surface, and
a coolant supply system for supplying coolant to the cooling channels, wherein the coolant supply system diverts a portion of a flow of coolant for internally cooling the aerofoils towards the cooling channels, the system comprising a plurality of passages extending across the front and/or the rear faces of the disc post from a base of the fixtures for the turbine blades to the forward and/or rearward ends of the cooling channels respectively.
2. A cooling arrangement according to
3. A cooling arrangement according to
4. A cooling arrangement according to
5. A cooling arrangement according to
7. A stage of the gas turbine engine comprising the turbine disc, the row of turbine blades attached to the turbine disc, and configured with the cooling arrangement for the turbine disc according to
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The present invention relates to a cooling arrangement for a turbine disc of a gas turbine engine
With reference to
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
Internal convection and external films are the prime methods of cooling the gas path components—airfoils, platforms, shrouds and shroud segments etc. The NGVs 31 consume the greatest amount of cooling air on high temperature engines. The HP blades 32 typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
The HP turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
A front cover-plate 38 covers the outer front face of the disc 33 and a rear cover-plate 39 covers the outer rear face of the disc. The cover-plates protect the outer front and rear faces of the disc and secure the blades on the disc. The front plate in particular also assists in directing a flow A of cooling air bled from the high pressure compressor to the disc and another flow B to the roots of the blades 32 (and thence into a cooling flow C for the interior of the blades). A row of aft lock-plates 40 also help to prevent axial movement of the blades relative to the disc.
The rotating structures, such as the disc 33, cover-plates 38, 39 and lock-plates 40 etc. are particularly sensitive to small changes in metal temperature, and rely heavily on copious amounts of cooling air to maintain their structural integrity. The prime cause of premature failure or more correctly part replacement for these critical components is low cycle mechanical and thermal fatigue. In recent years the outer peripheral temperature of the disc, and particularly the disc posts 35, has become a life-limiting location of the HP turbine.
Thus one proposal, illustrated in
Problems with this arrangement, however, are the relatively low heat transfer coefficients on the exposed outer surface 42 of the disc post 35, the lack of control in terms of flow direction in the cavity 41, and the mixing of the fresh and spent coolant within the cavity, which elevates the coolant temperature. In addition, heat radiates from the hot platforms 36 onto the exposed outer surface 42.
U.S. Pat. No. 7,207,776 (hereby incorporated by reference) proposes a cooling arrangement similar to the one shown in
Presently, disc life can be limited by the temperature of the disc post at its exposed outer surface even with the benefits of the cooling arrangements shown in
Thus there is a need for improved cooling arrangements for the disc periphery, and in particular for the exposed outer surface of the disc post.
Accordingly, a first aspect of the present invention provides a cooling arrangement for a turbine disc of a gas turbine engine, the turbine disc having a plurality of circumferentially spaced disc posts forming fixtures therebetween for a row of turbine blades, each turbine blade having an attachment formation which engages at a respective fixture, a platform radially outwardly of the attachment formation such that the adjacent platforms of the row form an inner endwall for the working gas annulus of the engine, and an aerofoil which extends radially outwardly from the platform, wherein a respective cavity is formed between an exposed radially outer surface of each disc post and the inner endwall;
wherein the cooling arrangement comprises:
at each disc post, a cooling plate removably locatable in the respective cavity and spaced radially outwardly from the exposed outer surface of the disc post to form a cooling channel between the cooling plate and the exposed outer surface, each cooling plate covering substantially the entire exposed outer surface, and
a coolant supply system for supplying coolant to the cooling channels.
The cooling channel allows the flow coolant therein to remain attached to the exposed outer surface of the disc post, and can help to prevent mixing between fresh and spent coolant. In this way, heat transfer can be enhanced and relatively high heat transfer coefficients obtained across most, if not all, of the exposed surface. Further, the cooling plate can protect the exposed surface from heat radiated across the cavity from the endwall. Although, the cooling plate is then itself subject to this radiated heat, being removably locatable, the cooling plate can be removed and replaced when necessary, e.g. at overhaul, thereby helping to increase the life of the more costly and critical disc.
The cooling arrangement may have any one or, to the extent that they are compatible, any combination of the following optional features.
Typically, each turbine blade has a fir tree attachment formation, the fixtures formed between the disc posts providing correspondingly-shaped recesses which co-operatively engage with the formations.
Preferably, each cooling plate has one or more exit holes formed therein for transferring spent coolant from the cooling channel to the cavity. In this way, the spent coolant can be prevented from mixing with fresh coolant entering the channel. The spent coolant can exit into the working gas annulus from the cavity by leakage between adjacent platforms or other pathways.
Preferably, the or each exit hole is positioned axially centrally in the cooling plate, and the coolant supply system supplies coolant to both the forward and the rearward ends of the cooling channel. This configuration allows both the forward and rearward edges of the exposed outer surface of the disc post to experience enhanced cooling, these edges generally experiencing a high heat load. However, alternatively, the or each exit hole can be positioned toward one of the forward and rearward edges of the exposed outer surface and the coolant supply system can supply coolant to the opposing end cooling channel. This allows a simpler and cheaper coolant supply system to be provided.
Conveniently, the coolant supply system may divert a portion of a flow of coolant for internally cooling the aerofoils towards the cooling channels, the system comprising a plurality of passages extending across the front and/or the rear faces of the disc post from a base of the fixtures for the turbine blades to respectively the forward and/or rearward ends of the cooling channels.
At each disc post, support formations may be formed in the flanks of the neighbouring turbine blades, the support formations supporting the cooling plate to maintain the radial spacing between the cooling plate and the exposed outer surface. For example, each support formation can be an additional tooth of a fir tree attachment formation. The formations can maintain the spacing of the cooling plate to the exposed outer surface against centrifugal forces acting on the plate. Alternatively, or additionally, lock-plates can be provided at the fore and aft faces of the disc posts to prevent axial movement of the blades relative to the disc, and, at each disc post, support formations can be formed in the respective fore and aft lock-plates, the support formations likewise supporting the cooling plate to maintain the radial spacing between the cooling plate and the exposed outer surface.
A separating member may be positioned in the cooling channel between the cooling plate and the exposed outer surface, the separating member containing a plurality of holes and dividing the cooling channel into a radially outer channel and a radially inner channel. The coolant supply system can then supply the coolant to the outer channel, whereupon the supplied coolant enters the inner channel via the holes as a series of jets which impinge on the exposed outer surface. By forming such impingement jets, enhanced cooling of the exposed outer surface can be obtained.
A second aspect of the present invention provides a cooling plate for the cooling arrangement of the first aspect.
A third aspect of the present invention provides the combination of a cooling plate of the first aspect and the optional separating member of that aspect.
A fourth aspect of the present invention provides a turbine blade for the cooling arrangement of the first aspect when the optional support formations are formed in the flanks of the neighbouring turbine blades, the support formations supporting the cooling plate to maintain the radial spacing between the cooling plate and the exposed outer surface.
A fifth aspect of the present invention provides a stage of a gas turbine engine comprising a turbine disc, a row of turbine blades attached to the turbine disc, and a cooling arrangement for the turbine disc according to the first aspect.
A sixth aspect of the present invention provides a gas turbine engine having the stage of the fifth aspect.
Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:
The present invention aims at improving the effectiveness of cooling arrangements for the disc post feature of particularly an HP turbine disc by: (i) elevating the heat transfer coefficients on the exposed outer surfaces of the disc posts and providing more control over the heat transfer coefficient distribution, (ii) helping to prevent fresh coolant from mixing with spent heated coolant, and (iii) reducing radiative heating of the disc posts by the high temperature of the underside of the blade platforms.
A flow B of pre-swirled air is fed through a series of holes in the front cover-plate 38, up the front face of the disc 33 and into the bucket grooves 60 that feed the entrances to the internal cooling system of the turbine blades. A portion D′ of this air is diverted off and passes radially outwards through a series of grooves machined into the front and rear faces of the disc posts 35 adjacent to the firtree attachment formations 34. Alternatively the grooves can be machined into thickened front 44 and rear 40 lock-plates instead of the disc posts. At each disc post, these front and rear streams of fresh coolant then turn at right angles towards one another through a radially narrow channel 51 formed between the exposed outer surface 42 of the disc post and a removably replaceable plate 50 that extends parallel to and covers substantially the entirety of the exposed outer surface. The plate is maintained in position against centrifugal forces by support projections 53 formed on the inner surfaces on the front and rear lock-plates, although the plate can be free to move in the gap between the outer surface 42 and the support projections 53 when the engine is stationary.
The coolant flow D′ over the exposed outer surface 42 is guided towards the centre of the disc post 35, where it exits the channel 51 through a hole or series of holes 52 in the plate 50 into the main part of the cavity 41 beneath the blade platforms 36. This flow is then re-used to cool the under platform region before being ejected back into the gas-path through slots in the dampers of the platforms or through the gaps between the platforms.
Preferably, the plate 50 has a relatively close dimensional fit to the surrounding surfaces of the firtree attachment formations 34 and the lock-plates 40, 44 to prevent excessive leakage around the plate.
The close proximity of the plate 50 to the exposed outer surface 42 of the disc post 35 helps to ensure a high velocity and a high Reynolds number coolant flow, which in turn can produce high heat transfer coefficients on the outer surface.
The guidance to the coolant flow D′ provided by the plate 50 also helps to ensure that fresh coolant entering the channel 51 is not diluted with spent coolant in the main part of the cavity 51 outboard of the plate.
The plate 50 provides a thermal barrier protecting the exposed outer surface 42 from the heat radiating from the lower surfaces of the blade platforms 36. Further, the plate can be readily replaced at overhaul if it is damaged by excessive exposure to high temperatures.
Although in
Another option is to introduce flow interrupting features such as trip strips, pin, fins, pedestals which project from the lower surface of the plate 50 in order to promote turbulence and hence mixing of coolant in the channel 51. In particular, trip strips can help to limit boundary layer thickening on the lower surface of the plate by causing boundary layer separation and reattachment. This may be beneficial if the plate is exposed to high temperatures by radiation from the underside of the platforms 36.
The second embodiment is identical to the first embodiment except that the plate 50 is maintained in position against centrifugal forces by axially-extending support projections 54 formed on the flanks of the neighbouring fir tree attachment formations 34. Each of these projections is effectively an additional tooth on the fir tree, and is thus relatively straightforward to manufacture at little additional cost during the grinding operation that is typically used to produce the fir tree serrations. Each projection can be axially continuous or composed of axially spaced portions (in
The arrangements of the third and fourth embodiments shown in
The cooling arrangement of the fifth embodiment is similar to that of the first embodiment except that a separating plate 55 between the plate 50 and the outer surface 42 forms an outer 51a and an inner 51b channel. The separating plate contains a pattern of through-holes 56. The cooling flow D′ is supplied as before up the front and rear faces of the disc post 35, and is then bled respectively rearward and forward into the ends of the outer channel 51a. This channel acts as a plenum chamber for impingement jets that are formed by the coolant as it passes through the through-holes 56 on its way from the outer to the inner channel. The jets are directed radially inwards onto the outer surface 42 to enhance the cooling of that surface. The flow that exhausts from the impingement jets then travels along the inner channel towards the centre of the disc post and exits into the main part of the cavity 41 through the one or more central holes 52.
Providing that the pressure ratio across the impingement jets is not too low, the convective cooling provided by impingement jets should be superior to that produced by the simpler arrangements of the previous embodiments.
As in the previous embodiments, the impingement arrangement can be supplied by only a front or rear coolant flow, or both as shown in
The heat transfer distribution across the exposed outer surface 42 can be controlled by varying the diameters, shapes and/or the distribution of the through-holes 56. However, cross flow in the inner channel 51b will tend to produce a distribution that provides greater rates of heat transfer at locations distant from the hole or holes 52. Thus, in the example shown in
A drawback associated with arrangement of the fifth embodiment is the additional weight of the separating plate 55 and a higher cost of manufacture.
Thus the present invention can provide the following advantages:
While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.
Tibbott, Ian, Jackson, Dougal R., Clarkson, Rory J.
Patent | Priority | Assignee | Title |
10167722, | Sep 12 2013 | RTX CORPORATION | Disk outer rim seal |
10619490, | Dec 19 2016 | Rolls-Royce Deutschland Ltd & Co KG | Turbine rotor blade arrangement for a gas turbine and method for the provision of sealing air in a turbine rotor blade arrangement |
10787920, | Oct 12 2016 | General Electric Company | Turbine engine inducer assembly |
11098593, | May 18 2018 | MTU AERO ENGINES AG | Rotor blade for a turbomachine |
11162366, | Feb 19 2019 | SAFRAN AIRCRAFT ENGINES | Rotor disc with axial stop of the blades, assembly of a disc and a ring and turbomachine |
11466582, | Oct 12 2016 | General Electric Company | Turbine engine inducer assembly |
11486252, | Sep 04 2018 | SAFRAN AIRCRAFT ENGINES | Rotor disc with axial retention of the blades, assembly of a disc and a ring, and turbomachine |
11591916, | Jul 02 2021 | Hamilton Sundstrand Corporation | Radial turbine rotor with complex cooling channels and method of making same |
11732592, | Aug 23 2021 | General Electric Company | Method of cooling a turbine blade |
11846209, | Oct 12 2016 | General Electric Company | Turbine engine inducer assembly |
Patent | Priority | Assignee | Title |
4505640, | Dec 13 1983 | United Technologies Corporation; UNITED TECHNOLOGIES CORPORATION, A DE CORP | Seal means for a blade attachment slot of a rotor assembly |
4505642, | Oct 24 1983 | United Technologies Corporation | Rotor blade interplatform seal |
5281097, | Nov 20 1992 | General Electric Company | Thermal control damper for turbine rotors |
5388962, | Oct 15 1993 | General Electric Company | Turbine rotor disk post cooling system |
5957660, | Feb 13 1997 | Rolls-Royce Deutschland Ltd & Co KG | Turbine rotor disk |
6017189, | Jan 30 1997 | SAFRAN AIRCRAFT ENGINES | Cooling system for turbine blade platforms |
6457935, | Jun 20 2001 | SAFRAN AIRCRAFT ENGINES | System for ventilating a pair of juxtaposed vane platforms |
7214034, | May 30 2002 | SAFRAN AIRCRAFT ENGINES | Control of leak zone under blade platform |
20040115054, | |||
20050201857, | |||
20050232751, | |||
20070041836, | |||
EP502660, | |||
EP1028228, | |||
EP1251243, | |||
GB2411697, | |||
GB2435909, | |||
WO2005095761, |
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